In a relatively modern aero-gas turbine engine certain accessories, such as the gearbox and electrical starter/generator, are mounted outwardly of the fan casing within the nacelle or airframe within which the engine is embedded. Accessory services, such as oil feed pipes and electrical cables, are routed through fairings spanning across the bypass duct. These fairings do not carry structural loads, but provide an aerodynamic shape around the accessory services.
To minimise aerodynamic drag the nacelle or airframe is tightly wrapped around the engine, minimising frontal area. However, one disadvantage is that aerodynamic shape of the airframe or nacelle is compromised with bulging to accommodate the accessories. Clearly, any bulging can be streamlined but by implication will be detrimental to the aircraft drag coefficient due to steeper cowl angles required to clear the engine accessories. For super-sonic aircraft, such bulging is also known to increase the sonic boom of the nacelle.
GB744,695 discloses a compact two-circuit gas turbine engine comprising a core engine having, in downstream flow sequence, a compressor, a combustor and a turbine. The core gas flow is turned and directed forwardly to flow through the combustor which is housed in an array of discrete tubes. The engine further comprises discrete bypass flow tubes which are, in a circumferential direction, alternately spaced between the combustor tubes. As the combustor tubes extend only an axial portion of the bypass tubes, engine accessories are housed between the bypass tubes and axially forward of the combustor tubes. Although this engine configuration is shorter by virtue of the reverse flow combustor, it is seriously disadvantaged as reversing the gas flow induces substantial flow energy losses and causes gas flow disruption into the combustor. Further, the circumferentially alternating bypass and combustor tubes mean that for any given air flow through the core engine, not only is there an annular inlet, but also the reversed combustor core flow occupies a substantial portion of what would be, in a modern conventional gas turbine engine, a substantially annular bypass duct. Thus the bypass gas flow is subject to substantial energy losses ingressing, flowing through and egressing the discrete bypass tubes. Thus the frontal area of this engine would be significantly greater than a conventional gas turbine engine having an annular bypass duct and no reversed combustor flow. Furthermore, GB744,695 does not disclose either a fairing spanning across a bypass duct or of mounting accessories within such a fairing. The engine of GB744,695 is not suitable for high-speed or super-sonic flight.